This invention relates generally to turbine engines and, more particularly, to methods and apparatus for cooling turbine engine blades and blade platforms.
High pressure turbine blades include an airfoil that is prone to trailing edge root cracks. Propagation of these cracks leads to eventual liberation of the airfoil. The cracks can potentially progress to a complete corn-cobbed rotor. The cracks are caused, at least in part, by blade components experiencing gas temperatures beyond the material capabilities.
To satisfy blade life requirements, the airfoils typically are cooled during operation. Airfoil cooling typically is achieved by convection cooling, e.g., in serpentine passages and film openings, and by film cooling which provides a protective layer of relatively cool air over an external surface of the airfoil. Cooling requirements are typically set by high temperature component life requirements for creep rupture and oxidation at the turbine blade operating conditions.
Cracking may be aggravated by skewed dovetails and sharp pressure side bleed slot geometric configurations for the blades. These configurations may cause very early trailing edge root crack indications in factory test engines.
For example, in the art of turbine blade cooling, it is well known to align the openings in the airfoil and the platform with airfoil regions experiencing high flow path gas temperatures. Generally, thermal gradients within a given radial span, i.e., low thermal gradient between blade bulk and its edges, are reduced. Additionally, cooling levels are matched with the mechanical stresses experienced in the rotating environment.
Accordingly, it would be desirable to provide a cooling configuration that improves cooling near the root trailing edge. It would be further desirable to reduce thermal stresses in a given radial span, in particular at the trailing edge region. It would be still further desirable if the reduced thermal stresses in the trailing edge vicinity prolonged low cycle fatigue life of the blades.